The subject matter disclosed herein is generally directed toward systems and methods for bringing a spacecraft to a desired attitude without using a gyroscope and, more particularly, toward systems and methods for estimating spacecraft angular rate information using star position measurements.
A spacecraft in order to perform its mission generally requires precise control of the orientation of the spacecraft payload relative to its target. A major challenge for vehicle designers is “pointing performance”, that is, making sure that the payload, for example an antenna, is aimed at the right spot or target on the Earth or elsewhere. Pointing is often autonomous, meaning directed by onboard computers, as opposed to having a team on the ground constantly commanding the vehicle.
There are two major components affecting spacecraft autonomous pointing performance. One component is knowledge of attitude estimation called estimator. Attitude estimation is to know the satellite's orientation with respect to some reference frame. The other major component is attitude controller. The controller defines ways to stabilize the system, and correct the satellite's attitude based on the command attitude versus the estimated attitude from the attitude estimator
Spacecraft attitude control system requires an on-board attitude determination subsystem and attitude controller. The logic of the attitude controller applies a torque command to the spacecraft to correct the error between a desired attitude and an estimated attitude. Spacecraft angular rate information is fed back for attitude control. In general, the body angular rate information is provided by three-axis gyros incorporated in inertial reference units (IRUs). A satellite typically includes at least two IRUs for redundancy in case one fails. The IRUs measure or determine the angular velocity about each of the three orthogonal axes. Angular velocity may be defined as the spin rate of the vehicle about each one of the three axes. Each IRU typically include three gyroscopes or simply gyros. Each gyro measures angular velocity about a different axis. Some IRUs are internally redundant, i.e., they include four gyros. To decrease the cost of a spacecraft, the number of gyros can be reduced, but this in turn may result in an increased risk of mission failure.
Due to limited resources or gyro failure, spacecraft attitude control systems can face drastic challenges in the area of system stability and fault autonomy. Various designs have been proposed to address satellite attitude control without a gyroscope. For example, the information needed to re-orient a disoriented spacecraft in three-dimensional space may be provided by a star tracker (comprising, e.g., imaging hardware and a high-speed computer processor).
It is known to use star trackers to determine spacecraft attitudes. But it is not common practice to use a star tracker to estimate the spacecraft rates in the absence of gyros. Estimated rates from star trackers can compensate for degraded mission performance in the presence of gyro failures or can be used as a reliable rate reference for sensor calibration.
Current star trackers use active pixel sensor-based chips and incorporate fast processors to produce derived rate measurements together with attitude information. But without proper rate estimate algorithms like the ones presented below, their rate estimates are highly unreliable.
A method for determining spacecraft angular velocity directly from star tracker measurements was disclosed by John L. Crassidis in a paper entitled “Angular Velocity Determination Directly from Star Tracker Measurements” AIAA Journal of Guidance, Control, and Dynamics, Vol. 25, No. 6, November-December 2002, pp. 1165-1168, and referenced hereinafter as Crassidis. Crassidis teaches a method for determining spacecraft angular velocity using a least-squares approach based only on knowledge of the star position vectors in the spacecraft body frame, which are obtained directly from the star tracker. Crassidis state the main advantage of his methodology is that it requires no information about the star reference vectors or the spacecraft attitude.
Singla et al. disclosed Kalman filtering algorithms for gyroless estimation of spacecraft body angular rates using a star tracker in a paper entitled “Spacecraft Angular Rate Estimation Algorithms for Star Tracker-Based Attitude Determination”, 13th AAS/AIAA Space Flight Mechanics Meeting, Ponce, Puerto Rico, February 2003, AAS Paper #03-191, and referenced hereinafter as Singla. Singla disclosed two embodiments. In a first embodiment, body angular rates are estimated with the spacecraft attitude using a dynamical model of the spacecraft and star identification. A second embodiment is capable of estimating the body angular rates independent of spacecraft attitude. The second embodiment employs a rapid update rate of a star tracker camera and finite difference analysis of “image flow” trajectories of the measured star line-of-sight vectors in the star tracker coordinate system.